Mechanisms for deploying and actuating airfoil-shaped bodies on unmanned aerial vehicles

ABSTRACT

Deployment and control actuation mechanisms are incorporated in unmanned aerial vehicles having folding wings and/or folding canards and/or a folding vertical stabilizer. The folding canards and folding vertical stabilizer can be deployed using respective four-bar over-center mechanisms. Elevators pivotably mounted to the folding canards and a rudder pivotably mounted to the folding vertical stabilizer can be controlled by means of respective twist link mechanisms. The folding wings have respective wing roots that are driven by respective gas springs to pivot on bearings about a wing root hub having control servo wire paths.

BACKGROUND

This disclosure generally relates to mechanisms for deploying andactuating airfoil-shaped bodies on compressed carriage unmanned aerialvehicles. In particular, this disclosure relates to mechanisms fordeploying wings, canards or vertical stabilizers and mechanisms foractuating trailing edge control surfaces on compressed carriage unmannedaerial vehicles.

Unmanned surveillance air vehicles, glide munitions, winged missiles,and other types of unmanned aerial vehicles (UAVs) are sometimesconfigured to be carried internally or externally on a larger motheraircraft (or a submarine). Because the carried UAV itself is usuallysmall and typically has a limited range, it is flown to a location nearto where it is to perform its mission, as cargo on the mother aircraft,and then air launched to perform the mission. The carried UAV may laterbe recovered, or it may be considered expendable and destroyed at thecompletion of the mission.

A known carried UAV has laterally extending wings and canards and avertical stabilizer which, in the absence of foldability, would make itawkward to store the carried UAV on the mother aircraft or submarine. Tofacilitate the internal or external storage and transport on the motheraircraft or submarine, the carried UAV may be provided with foldingwings, canards and vertical stabilizer. The folding wings, canards andvertical stabilizer are in storage positions during carriage, and thenare unfolded to a deployed flight position shortly after launch from themother aircraft or submarine. The UAV is launched from a mother aircraftat high speed in a compressed carriage state. First a balloon isinflated to slow down the UAV and then a parachute is deployed tostabilize the UAV. The folded wings, canards and vertical stabilizer arethen deployed and the engine of the UAV is started. The UAV is then ableto fly under remote control by the mother aircraft. The UAV could alsobe controlled by a remote ground control station, ship, submarine, etc.

There is a need for improved mechanisms for deploying folding wings,canards and vertical stabilizers and actuating control surfaces on UAVsof the foregoing type.

SUMMARY

The deployment and control actuation mechanisms disclosed herein can beincorporated in UAVs having folding wings and/or folding canards and/ora folding vertical stabilizer. In accordance with various aspects of thedisclosed subject matter, the folding canards and folding verticalstabilizer can be deployed using respective four-bar over-centermechanisms. In accordance with other aspects of the disclosed subjectmatter, elevators pivotably mounted to the folding canards and a rudderpivotably mounted to the folding vertical stabilizer can be controlledby means of respective twist link mechanisms. In accordance with furtheraspects, the folding wings have respective wing roots that are driven byrespective linear actuators to pivot on bearings about a wing root hubhaving control servo wire paths.

The folding wing mechanism comprises a low-profile, low-friction,self-powered high-load-capacity wing opening mechanism and a compactwing lock mechanism that permit a small air/submarine-launched UAV tocompress into a small package by folding and locking the wings beneaththe UAV fuselage. Once the UAV is launched, a locking mechanism forlocking the wings in the stowed condition is unlocked and then linearactuators drive the wings open independently and passively. The lockingmechanism then locks the wings in the deployed condition. The mechanismis strong enough to support air loads and wing hook recovery loads. Thismechanism permits folding of a small UAV wing assembly for compressedcarriage and subsequent air- or sub-launched deployment. It permits asmall UAVs wings to compress into an efficient packaging scheme forfitment into a small launch container, rapid deployment, and high loadaircraft recovery via rope capture by a wing mounted hook. The mechanismalso permits a path for the aileron/flap control servo wires to passthrough the mechanism and into the wings to the control servos in thefolded and deployed conditions. The wing deployment mechanism, lockmechanism, and deployment actuator are all self contained within thewing root.

The folding canard mechanism comprises a spring-driven four-barover-center mechanism for canard deployment, that employs a twist linkmechanism for elevator operation. The combination of these twomechanisms allows for rapid deployment of a compressed carriage smalltactical UAV with quarter-chord elevator operation. These mechanismspermit folding of a small UAV canard/elevator for compressed carriageand subsequent air- or sub-launched deployment. They permit a small UAVwith canards to compress into an efficient packaging scheme for fitmentinto a small launch container, rapid deployment and elevator controlsurface operation once uncompressed.

The folding vertical stabilizer mechanism provides means for folding acompressed carriage UAV vertical stabilizer, including the verticalstabilizer deployment mechanism, stowed and deployed locking mechanismsand rudder control actuation mechanisms. These mechanisms permit foldingof a small UAV vertical stabilizer for compressed carriage andsubsequent air- or submarine-launched deployment. They permit a smallUAV to compress into an efficient packaging scheme for rapid deploymentand rudder operation once uncompressed.

In view of the foregoing, one aspect of the subject matter disclosedherein is an unmanned aerial vehicle comprising a fuselage, a deploymentmechanism supported by the fuselage, and a folding airfoil-shaped bodyattached to the deployment mechanism, wherein the deployment mechanismcomprises first and second links which are pivotably coupled to eachother, the folding airfoil-shaped body being in a stowed position whenthe first and second links are not aligned with each other and thefolding airfoil-shaped body being in a deployed position when the firstand second links are aligned with each other.

Another aspect of the subject matter disclosed herein is an unmannedaerial vehicle comprising a fuselage, a first airfoil-shaped body thatis rotatable relative to the fuselage from a stowed position to adeployed position, an actuation mechanism supported by the fuselage, anda second airfoil-shaped body pivotably coupled to a trailing edge of thefirst airfoil-shaped body, wherein the actuation mechanism comprises amotor, a first arm coupled to the motor, a second arm coupled to thesecond airfoil-shaped body, and a twist link that couples the first andsecond arms to each other, an angle of the second airfoil-shaped bodyrelative to the first airfoil-shaped body being adjustable in responseto movement of the first arm.

A further aspect of the subject matter disclosed herein is an unmannedaerial vehicle comprising a fuselage, a deployment mechanism supportedby the fuselage, and a wing attached to the deployment mechanism,wherein the deployment mechanism comprises a wing root hub attached tothe fuselage, a wing root pivotably coupled to the wing root hub, and anactuator pivotably coupled to the wing root hub and to the wing root atopposite ends thereof, wherein the wing is attached to the wing root androtates from a stowed position to a deployed position in response toextension of the actuator.

BRIEF DESCRIPTION OF THE DRAWINGS

Various embodiments will be hereinafter described with reference to thedrawings for the purpose of illustrating the foregoing and other aspectsof the disclosed subject matter.

FIGS. 1 and 2 are diagrams showing isometric views of a UAV havingfolding wings, folding canards and a folding vertical stabilizer indeployed and stowed positions respectively in accordance with oneembodiment.

FIG. 3 is a diagram showing an isometric view of a portion of a verticalstabilizer/rudder that has been removed from the vehicle.

FIGS. 4 and 5 are diagrams showing isometric views of a verticalstabilizer deployment mechanism comprising a four-bar over-centermechanism shown in a vertical stabilizer stowed state in FIG. 4 and in avertical stabilizer deployed state in FIG. 5.

FIG. 6 is a diagram showing the relationship of the attached verticalstabilizer/rudder to the deployment mechanism depicted in FIGS. 4 and 5.The fuselage has been removed.

FIG. 7 is a diagram showing an isometric view of a rudder controlactuation mechanism comprising a twist link with mono-ball and clevisrod ends, which permits rudder control in unison with verticalstabilizer deployment. A portion of a frame has been removed to showportions of a twist link that would otherwise be hidden.

FIG. 8 is a diagram showing an isometric view of a canard/elevatorassembly in accordance with one embodiment.

FIG. 9 is a diagram showing an isometric view of a canard deploymentmechanism comprising a four-bar over-center mechanism and deploymentsprings which reside internal to the canard aerodynamic surface.

FIGS. 10-12 are diagrams showing isometric views of a canard deploymentmechanism comprising a four-bar over-center mechanism shown in a canardstowed state in FIG. 10, in an intermediate position in FIG. 11 and in acanard deployed state in FIG. 12.

FIG. 13 is a diagram showing an isometric view of an elevator controlactuation mechanism comprising a twist-link with mono-ball rod ends,which permits ¼-chord elevator control in unison with canard deployment.

FIG. 14 is a diagram showing an isometric view of folded wings inisolation.

FIG. 15 is a diagram showing a sectional view of a wing deploymentmechanism comprising wing roots driven by linear actuators to pivot onx-type bearings about a wing root hub having control servo wire paths.

FIGS. 16 and 17 are diagrams showing sectional views of the wingdeployment mechanism with control servo wire path in a wing deployedstate in FIG. 16 and in a wing stowed state in FIG. 17.

Reference will hereinafter be made to the drawings in which similarelements in different drawings bear the same reference numerals.

DETAILED DESCRIPTION

FIGS. 1 and 2 show a UAV 2 comprising a fuselage 4, a pair of foldingwings 6, a pair of folding canards 8, a folding vertical stabilizer 10,and a propeller 12. FIG. 1 shows the UAV with the folding wings, canardsand vertical stabilizer deployed; FIG. 2 shows the UAV with the foldingwings, canards and vertical stabilizer folded, i.e., stowed forcompressed carriage. Each of the folding airfoil-shaped bodies has arespective deployment mechanism not shown in FIGS. 1 and 2. The verticalstabilizer 10 has a rudder 34 pivotably mounted to its trailing edge.Each canard 8 has an elevator 70 coupled to its trailing edge. Each wing6 also has an inboard flap 110 and an outboard aileron 114 pivotablycoupled to the trailing edge of the wing.

The deployment mechanism for the folding vertical stabilizer 10 (whichis described in detail below with reference to FIGS. 4-6) causes thestowed vertical stabilizer shown in FIG. 2 to pivot about an axis thatis generally normal to the stowed vertical stabilizer, the verticalstabilizer rotation being stopped when the vertical stabilizer isvertical or nearly vertical, as seen in FIG. 1. The deployment mechanismfor each folding canard 8 (which is described in detail below withreference to FIGS. 9-12) causes the stowed canard 8 shown in FIG. 2 topivot about an axis that is generally parallel to the longitudinal axisof the fuselage 4 (and perpendicular to the vertical stabilizer pivotaxis), the canard rotation being stopped when the canards extendlaterally in a horizontal plane, as seen in FIG. 1. The deploymentmechanism for each folding wing 6 (which is described in detail belowwith reference to FIGS. 14-17) causes the stowed wings shown in FIG. 2to pivot about an axis that is generally perpendicular to the pivot axesof the vertical stabilizer and of the canards, the wing rotation beingstopped when the canards extend laterally or nearly laterally in ahorizontal plane, as seen in FIG. 1.

FIG. 3 shows a portion of a vertical stabilizer 10 that has been removedfrom the vehicle. A rudder 34 is pivotably coupled to the trailing edgeof the vertical stabilizer. A vertical stabilizer deployment mechanismcomprising a frame 16 is mounted on the inside of the fuselage 4 andsupports a rudder control servo 38. The vertical stabilizer deploymentmechanism comprises a hub shaft 22 that projects outside the fuselage 4and to which the root 134 of vertical stabilizer 10 is fastened by meansof screws 136. When the vertical stabilizer root 134 is fastened to hubshaft 22 and in a stowed position (see FIG. 2), the deployment mechanismcan be unlocked to cause the vertical stabilizer to rotate from thestowed position to a deployed position (see FIG. 1). Item 36 in FIG. 3is a rudder control input arm that has been removed from the vehicle.The position and function of rudder control input arm 36 will bedescribed later with reference to FIGS. 6 and 7.

FIGS. 4 and 5 show the vertical stabilizer deployment mechanism inaccordance with one embodiment. This vertical stabilizer deploymentmechanism comprises a four-bar (i.e., four- link) over-center mechanism14, which is shown in a vertical stabilizer stowed state in FIG. 4 andin a vertical stabilizer deployed state in FIG. 5. The four linksinclude a vertical stabilizer mechanism frame 16 (third link), avertical stabilizer input link 18 (first link) has one end pivotablycoupled to vertical stabilizer mechanism frame 16 by joint A (with theaid of washers 15), a connecting link 20 (second link) has one endpivotably coupled to a clevis end of vertical stabilizer input link 18by joint B, and a vertical stabilizer root hub (fourth link) comprisinga large-diameter vertical stabilizer root hub shaft 22 and an arm 26pivotably coupled to a clevis end of connecting link 20 by joint C. Thevertical stabilizer root hub shaft 22, which serves as the verticalstabilizer folding joint, is also pivotably coupled to verticalstabilizer mechanism frame 16 and fuselage 4 by a pair of pivot bushings32. The folding vertical stabilizer (not shown in FIGS. 4 and 5) isattached to the vertical stabilizer root hub shaft 22 and rotates as thevertical stabilizer root hub rotates.

The over-center mechanism 14 is unlocked by activation of an unlockservo (not shown in the drawings) that controls a vertical stabilizerlock arm (not shown in the drawings) to disengage from a lock hook 21 onconnecting link 20 (see FIG. 4). The over-center mechanism 14, whenunlocked, is driven by a torsion spring 24, one end of which bearsagainst a protruding portion (not visible in FIG. 4) of joint B disposedbehind vertical stabilizer input link 18. The torsion spring 24 drivesthe over-center mechanism 14 from a first state (shown in FIG. 4)whereat the folding vertical stabilizer is in a stowed position to asecond state (shown in FIG. 5) whereat the folding vertical stabilizeris in a deployed position. More specifically, the torsion spring 24causes vertical stabilizer input link 18 to rotate clockwise (in theview of FIG. 4), which clockwise motion is converted to counterclockwiserotation of the vertical stabilizer root hub (and attached verticalstabilizer) via the connecting link 20.

The over-center mechanism 14 and the deployment torsion spring 24 resideinternal to the fuselage (not shown). Once unlocked from the stowedposition, the torsion spring 24 drives mechanism 14 open, causingvertical stabilizer input link 18 to rotate and connecting link 20 torotate/translate. Links 18 and 20 only stop translating/rotating whenlink 18 impacts a stop bolt 28 (shown in FIG. 5), which is bolted to thefuselage (not shown in FIG. 5). Rubber tubing 30 is added to snub impactforces and absorb the kinetic energy of links 18 and 20. As seen in FIG.5, when the over-center mechanism 14 locks itself in the verticalstabilizer deployed position, joints A, B and C are aligned. Onceover-center, no amount of force on the vertical stabilizer will permitthe mechanism to unlock.

The indirect drive of vertical stabilizer root hub shaft (output link)22 by torsion spring 24 on a four-bar input link allows the mechanism tolock in an over-center condition just as the vertical stabilizer reachesfull deployment. This combines the deployment mechanism and the deployedlocking mechanism into an efficient and compact design. Additionally,the inherent nature of a four-bar over-center mechanism is that theoutput link (the vertical stabilizer root hub) approaches zero velocityas the mechanism approaches the over-center condition. This prevents theoutput link from impacting a mechanical stop and prevents the largeimpulse loads of the direct drive system. The large-diameter verticalstabilizer root hub shaft 22 distributes the aerodynamic loads on thevertical stabilizer into the aircraft structure more efficiently, andreduces stress on the mechanism. Additionally, the large diameter allowshollowing of the root hub, which permits the rudder control mechanism tobe routed through the root hub and to the rudder.

FIGS. 6 and 7 shows a rudder control actuation mechanism comprising atwist link 88 which permits control of a rudder 34 (pivotably coupled tothe vertical stabilizer 10) in unison with vertical stabilizerdeployment. This is done by mounting the rudder control servo 38internal to the UAV fuselage with the servo crank arm 40 in plane with arudder control input arm 36 in the vertical stabilizer deployedcondition. The rudder control input arm 36 is installed within the root134 of the vertical stabilizer 10 and is fixedly connected to thepivotable rudder 34 by a rod 35. The twist link 88 connects the servocrank arm 40 to the rudder control input arm 36 and is concentric withthe vertical stabilizer root hub shaft centerline. This permits thevertical stabilizer 10 to fold and deploy without affecting the ruddercontrol actuation mechanism.

The twist link 88 comprises a threaded rod 92, a mono-ball rod end 90threadably coupled to one end of threaded rod 92 and a clevis rod end 94threadably coupled to the other end of threaded rod 92. The axis ofthreaded rod 92 lies along the vertical stabilizer root hub shaftcenterline. A distal end of servo crank arm 40 is coupled to themono-ball of mono-ball rod end 90 by a screw 42. The clevis rod end 94is pivotably coupled to one end of the rudder control input arm 36. Theangle of rudder 34 can be controlled by rudder control servo 38 by meansof the servo crank arm 40, the twist link 88 and the rudder controlinput arm 36, i.e., any rotation of servo crank arm 40 causes rudder 34to rotate about its pivot axis.

The mono-ball rod end 90 consists of a spherical ball with a holethrough it that fits into a socket that forms a ball-in-socket type ofjoint. The socket that the mono-ball fits into also contains femalethreads such that it can be attached to one end of threaded rod 92. Amono-ball rod ends is also commonly known as a “rod end bearing” or a“helm joint”. A mono-ball rod end is a joint that allows translation androtation of a link where one end is out of plane with the other end. Amono-ball rod end is limited in the amount of out-of-plane motion thatcan be made because the fastener that passes through the spherical ballwill eventually interfere with the sides of the socket. The ruddercontrol actuation mechanism shown in FIG. 7 employs a twist link 88 thatuses one female threaded mono-ball rod end 90 and one female threadedclevis rod end 94 screwed onto opposing ends of a length of threaded rod92. Only one rod end needs to be rotatable relative to the threaded rod92, so the other rod end can be locked to the threaded rod with a jamnut. This allows rod ends 90 and 94 to rotate relative to each otherabout the axis of threaded rod 92.

The axis of twist link 88 lies on the axis of rotation for thedeployment mechanism. When the deployment of the vertical stabilizeroccurs, the rudder control actuation mechanism twist link 88 simplyallows the two rod ends 90 and 94 to rotate about the axis of rotationof the deployment mechanism. In this manner, no “input” to rudder 34 ismade during deployment. (A slight input is made since the twist link 88grows slightly in length due to the threaded rod ends rotating, but thisinput is negligible.)

FIG. 8 shows a canard/elevator assembly in accordance with oneembodiment. The canard 8 is attached to a canard deployment mechanismcomprising a folding canard root 46 attached to the fuselage (notshown), a connecting link 48 pivotably coupled to folding canard root46, a pair of folding canard tumblers 50 pivotably coupled to connectinglink 48, and a folding mechanism frame 52 pivotably coupled to tumblers50. The canard deployment mechanism will be described below in moredetail with reference to FIGS. 9-12. The elevator 70 is pivotablycoupled to the trailing edge of the canard 8. The angle of the elevator70 is controlled by an elevator control actuation mechanism comprisingan elevator servo mounted to a canard servo bracket 64, which is alsoattached to the fuselage. The elevator control actuation mechanismcomprises a twist link that connects a servo crank arm 68 to an elevatorcontrol input arm 72. The twist link comprises a threaded rod 100, theends of which are respectively threadably coupled to first and secondmono-ball rod ends 98 and 102. One end of elevator control input arm 72is coupled to mono-ball rod end 98 by a screw 106, while one end ofservo crank arm 68 is coupled to mono-ball rod end 102 by a screw 104.The elevator control actuation mechanism will be described below in moredetail with reference to FIG. 13.

FIGS. 9-12 show the canard deployment mechanism in accordance with oneembodiment. Each canard has its own canard deployment mechanism. Thecanard deployment mechanism comprises a four-bar (i.e., four-link)over-center mechanism 44 driven by two tension springs 54 (see FIG. 9)mounted internal to the canard structure. This four-bar over-centermechanism 44 is shown in a canard stowed state in FIGS. 9 and 10, in anintermediate state in FIG. 11, and in a canard deployed state in FIG.12. The four links include a folding canard root 46 (third link), aconnecting link 48 (first link) pivotably coupled to folding canard root46 by a joint (not visible in the drawings), a pair of folding canardtumblers 50 (second link) pivotably coupled to connecting link 48 byrespective joints D (only one of which is visible in FIG. 11), and afolding mechanism frame 52 (fourth link) pivotably coupled to tumblers50 by respective joints E (only one of which is visible in FIG. 10). Thefolding mechanism frame 52 is also pivotably coupled to folding canardroot 46 by a piano hinge 62 (best seen in FIG. 10), which serves as thecanard folding joint. The folding canard (not shown in FIGS. 9-12) isattached to the folding mechanism frame 52 and deploys as the foldingmechanism frame 52 pivots.

The over-center mechanism 44 is unlocked by means of known stowed lockmechanisms with slider release to release a hook, which arrangement isnot shown in the drawings. This method is used to retain the foldingcanards in the compressed (stowed) condition until an unlock servo (notshown in the drawings) is signaled to release them.

Referring to FIG. 9, the over-center mechanism 44, when unlocked, isdriven to open by the pair of tension springs 54, the distal ends ofwhich are attached to a spring retainer 58 that is fixed relative to thecanard structure. The proximate ends of tension springs 54 are connectedvia respective spring extensions 56 to respective spring attachmentscrews 60 screwed into respective tumblers 50. The tension springs 54drive the over-center mechanism 44 from a first state (shown in FIG. 10)whereat the folding canard is in a stowed position to a second state(shown in FIG. 12) whereat the folding vertical stabilizer is in adeployed position. More specifically, the tension springs 54 causetumblers 50 to rotate counterclockwise (in the view of FIG. 10), whichcounterclockwise motion is converted to counterclockwise rotation ofconnecting link 48 and folding mechanism frame 52 (and attached canard).

The over-center mechanism 44 and the deployment tension springs 54reside internal to the canard aerodynamic surface (not shown in FIGS.9-12). Once unlocked from the stowed position, the tension springs 54drives mechanism 44 open, causing connecting link 48 to rotate andtumblers 50 to rotate/translate. Although not visible in FIG. 12, whenthe over-center mechanism 44 locks itself in the canard deployedposition, joints D and E and the joint where connecting link 48 ispivotably coupled to folding canard root 46 are aligned.

The canard hinges about a small-diameter piano hinge 62 (see FIG. 10).The piano hinge 62 resides internal to the fuselage profile such thatnone of the canard lifting surface is sacrificed to the hinge. Indirectdrive of the canard folding mechanism frame (output link) 52 by tensionsprings 54 pulling on folding canard tumblers 50 allows the mechanism tolock over-center just as the canard reaches full deployment. Thefour-bar over-center mechanism 44 combines the deployment mechanism andthe deployed lock mechanism into an efficient and compact design. Thiseliminates the need for a separate deployed lock mechanism.Additionally, the inherent nature of a four-bar over-center mechanism isthat the output link (folding mechanism frame 52 and canard) approacheszero velocity as the mechanism approaches the over-center condition.This prevents the canard mechanism from impacting a mechanical stop withthe inertia of the entire canard, and eliminates the need to increasethe size of the mechanism in order to handle the impact loads. Onceover-center, no amount of force on the canard will permit the mechanismto unlock. The spring force keeps the mechanism over-center in thecanard deployed position.

FIG. 13 shows an elevator control actuation mechanism comprising a twistlink 96 which permits control of a ¼-chord elevator 70 (pivotablycoupled to the canard) in unison with canard deployment. The twist link96 connects the servo crank arm 68 to the elevator control input arm 72.The twist link 96 comprises a threaded rod 100, a first mono-ball rodend 98 threadably coupled to one end of threaded rod 100 and a secondmono-ball rod end 102 threadably coupled to the other end of threadedrod 100. Only one mono-ball rod end needs to be rotatable relative tothe threaded rod 100, so the other mono-ball rod end can be locked tothe threaded rod with a jam nut. This allows mono-ball rod ends 98 and102 to rotate relative to each other about the axis of threaded rod 100.A distal end of servo crank arm 68 is coupled to the mono-ball of thesecond mono-ball rod end 102 by a screw 104. A proximal end of elevatorcontrol input arm 72 is coupled to the mono-ball of the first mono-ballrod end 98 by a screw 106. The angle of elevator 70 can be controlled byelevator control servo 66 by means of the servo crank arm 68, the twistlink 96 and the elevator control input arm 72, i.e., any rotation ofservo crank arm 68 causes elevator 70 to rotate about its pivot axis.

The twist link 96 permits rotation of the canard during deployment whilethe mono-ball rod ends allow the rudder control actuation mechanism tosweep an out-of-plane conic shape and maintain positive connection. Bymounting the control actuation servo mono-ball along the canardpiano-hinge line, little to no input to the elevator control surfaceoccurs during canard deployment, and the control linkage and servoresides entirely within the fuselage. This quarter-chord elevatorfolding canard technique is only possible because the canard piano hinge(item 62 in FIG. 10) resides internal to the fuselage profile.

The twist link 96 for the elevator 70 is employed in a similar manner asthe twist link for the rudder, but here only the input end of the twistlink lies on the axis of rotation of the deployment mechanism. So nowduring deployment, the control actuation twist link rod ends rotateapproximately 90 degrees relative to each other and also the rod sweepsout a conic. The placement of the input mono-ball on the axis ofrotation of the canard (or as close as possible) prevents anyuncommanded input to the elevator due to canard rotation duringdeployment. (Again a small input is made because the twist link grows inlength slightly, but this is negligible.)

FIG. 14 shows a removable wing assembly 108 in isolation, as compared toFIG. 2 which showed the same wing assembly installed on the fuselage.Each folded wing 6 also has a wing root 78 which is pivotably coupled toa wing root hub 76. Each folded wing 6 also has an inboard flap 110 andan outboard aileron 114 pivotably coupled to the trailing edge of thewing. Flap 110 and aileron 114 can be pivoted relative to the wing 6under the control of respective control servos 112 and 116 mountedinside the wing. The control servos 112 and 116 (as well as othercontrol servos previously mentioned) are controlled by an onboardcontroller (not shown in the drawings) situated inside the fuselage.Each control servo is connected to the controller by means of a controlservo wire 120 (shown in FIGS. 16 and 17). In accordance with theembodiment shown in FIGS. 15-17, the control servo wires 120 (shown inFIGS. 16 and 17) connecting the control servos 112 and 116 (shown inFIG. 14) on each wing to the onboard controller (not shown) pass throughrespective control servo wire openings 118 (shown in FIG. 15) in thewing root hub 76.

The sectional view of FIG. 15 shows one wing root 78 pivotably coupledto the upper half of wing root hub 76 by means of X-type bearings 80.The other wing root (not shown) is pivotably coupled to the lower halfof wing root hub 76. The wing root hub 76 is attached to the fuselage bymeans of six high-strength hub attachment bolts 82 (only four are seenin FIG. 15), which pass through holes in the wall of wing root hub 76and are inserted from outside of the fuselage. This permits the wingassembly to be removed from the fuselage without disassembly of the UAV.

FIGS. 16 and 17 show sectional views of the wing deployment mechanismwing deployed and wing stowed states respectively. The wing deploymentmechanism for each wing comprises a large-diameter wing root hub 76, awing root 78, and a low-profile linear actuator 84 that is mountedinternally to the respective wing root. The linear actuator is disposedin a passageway 74 formed in wing root 78. The control servo wire 120also passes through passageway 74 on its way from the inside of the wingto the interior space of the wing root hub 76.

Still referring to FIGS. 16 and 17, one end of linear actuator 84 ispivotably coupled to the wing root hub 76 by a joint 124 and the otherend of linear actuator 84 is pivotably coupled to the wing root 78 by ajoint 126. FIG. 16 shows the position of the linear actuator relative tothe wing root hub when the wing is deployed; FIG. 17 shows the positionof the linear actuator relative to the wing root hub when the wing isstowed. In a wing stowed (i.e., folded) position, the wing root 78 isprevented from rotating by a servo-controlled retractable lock pin (notshown in FIGS. 16 and 17) mounted inside the wing root hub 76 thatengages a first slot 128 (see FIG. 17) on the inner periphery of thewing root 78. When the lock pin is retracted by the unlock servo (notshown), the linear actuator 84 rotates the wing root to the deployedposition. A stop block 122 attached to the wing root hub 76 andextending radially outward prevents rotation of the wing beyond adesired limit. The retractable lock pin (which has a spring urging itradially outward) will ride the inner surface of the wing root until thewing root reaches the deployed condition and hits the stop block 122,where the spring will urge the lock pin (not shown) to engage a secondslot 130 (see FIG. 17) on the inner periphery of the wing root 78,thereby locking the wing root in the deployed position in a well-knownmanner.

Referring again to FIG. 15, the X-type ball bearings 80 are capable ofsupporting axial, thrust, and bending loads with very low friction andfree play. Since two ball bearings are installed per wing root, an openarea between the ball bearings exists in each wing root that permitsboth the linear actuators and control servo wires to pass between them(see space 132 in FIG. 15). This configuration solves load distribution,free play, and friction problems, as well as the control servo wirerouting problem and eliminates the need for a large gas spring withinthe fuselage or motor power from the UAV to drive the wings open.Additionally, the open space between each X-type ball bearing in thewing roots creates a space to put mechanical “hard stop” stop blocks 122and locking features in the wing roots.

In summary, when the wing unlock servo is actuated, the lock pinsretract permitting the wings to rotate. The linear actuators of thedeployment mechanism will then drive the wings open. If electricallypowered linear actuators were employed, then power from the UAV would berequired to actuate them. The same servo wire path that is used to runservo wires from the wing root to the wings could likewise be used toprovide power to the linear actuators.

While various embodiments have been described, it will be understood bythose skilled in the art that various changes may be made andequivalents may be substituted for elements thereof without departingfrom the scope of the teachings herein. In addition, many modificationsmay be made to adapt a particular situation to those teachings withoutdeparting from the essential scope thereof. Therefore it is intendedthat scope of the claims set forth hereinafter not be limited to thedisclosed embodiments.

The invention claimed is:
 1. An unmanned aerial vehicle comprising afuselage, a deployment mechanism supported by said fuselage, and anairfoil-shaped body attached to said deployment mechanism and rotatablebetween stowed and deployed positions, wherein said deployment mechanismcomprises: first and second links which are pivotably coupled to eachother at a first joint, a third link that is attached to said fuselageand pivotably coupled to said first link at a second joint, and a fourthlink comprising a first portion that is pivotably coupled to said secondlink at a third joint and a second portion that is pivotably coupled tosaid third link, said fourth link being rotatable about an axis ofrotation, wherein said airfoil-shaped body is attached to said secondportion of said fourth link and is not attached to any of said first,second and third links, and wherein said airfoil-shaped body can rotatewith said fourth link from said stowed position at which said first,second and third joints are not aligned to a deployed position at whichsaid first, second and third joints are aligned.
 2. The unmanned aerialvehicle recited in claim 1, wherein said folding airfoil-shaped bodycomprises a folding canard, further comprising a tension spring thaturges said first and second links from an unaligned state to an alignedstate.
 3. The unmanned aerial vehicle recited in claim 1, wherein saidairfoil-shaped body comprises a vertical stabilizer, further comprisinga torsion spring that urges said first and second links from anunaligned state to an aligned state.
 4. The unmanned aerial vehiclerecited in claim 1, wherein said fourth link is pivotable from a firstangular position whereat said fourth link does not abut said third linkto a second angular position whereat said fourth link abuts said thirdlink as said first and second links move from an unaligned state to analigned state.
 5. The unmanned aerial vehicle recited in claim 1,further comprising a stop bolt attached to said fuselage and a stop boltcovering made of elastomeric material, said stop bolt being positionedso that said first link presses against said stop bolt covering whensaid first and second links are aligned.
 6. The unmanned aerial vehiclerecited in claim 1, wherein said first joint is disposed between saidsecond and third joints when said first, second and third joints arealigned.
 7. The unmanned aerial vehicle recited in claim 1, wherein saidthird joint is disposed between said first and second joints when saidfirst, second and third joints are aligned.
 8. An unmanned aerialvehicle comprising a fuselage, a first airfoil-shaped body that isrotatably mounted to said fuselage for rotation from a stowed positionto a deployed position about a first axis of rotation, an actuationmechanism supported by said fuselage, and a second airfoil-shaped bodypivotably coupled to a trailing edge of said first airfoil-shaped bodyfor rotation about a second axis of rotation, wherein said actuationmechanism comprises a motor, a first arm coupled to said motor, a secondarm coupled to said second airfoil-shaped body, and a twist link thatcouples said first and second arms to each other, an angle of saidsecond airfoil-shaped body relative to said first airfoil-shaped bodybeing adjustable in response to movement of said first arm.
 9. Theunmanned aerial vehicle recited in claim 8, wherein said twist linkcomprises a threaded rod and first and second rod ends coupled to saidthreaded rod in a manner such that said first and second rod ends arerotatable relative to each other about an axis of said threaded rod,wherein said first rod end comprises a first ball attached to said firstarm and a first socket supporting said first ball.
 10. The unmannedaerial vehicle recited in claim 9, wherein said first airfoil-shapedbody is a vertical stabilizer and said second airfoil-shaped body is arudder pivotably coupled to said vertical stabilizer.
 11. The unmannedaerial vehicle recited in claim 10, wherein said second rod endcomprises a clevis pivotably coupled to said second arm.
 12. Theunmanned aerial vehicle recited in claim 9, wherein said firstairfoil-shaped body is a canard and said second airfoil-shaped body isan elevator pivotably coupled to said canard.
 13. The unmanned aerialvehicle recited in claim 12, wherein said second rod end comprises asecond ball attached to said other end of said second arm and a secondsocket supporting said second ball.
 14. The unmanned aerial vehiclerecited in claim 9, wherein one of said first and second rod ends isrotatable relative to said threaded rod.
 15. The unmanned aerial vehiclerecited in claim 9, further comprising a deployment mechanism supportedby said fuselage, said deployment mechanism comprising a hinge, saidfirst airfoil-shaped body being attached to said deployment mechanism,and said first ball being disposed near a hinge line of said hinge.